1. Field of the Invention
The invention relates to aircraft environmental control systems powered by gas turbine engines for supplying aircraft system bleed air. In particular, the invention relates to an aircraft gas turbine engine powered bleed air supply system for an aircraft environmental control system wherein the excess pressure energy is returned to the engine to improve fuel efficiency and the system is also used to reduce drag on engine nacelles or other surfaces by pumping boundary layer air.
2. Description of Related Art
Environmental control systems, commonly referred to as ECS systems, incorporate various pieces of equipment such as turbocompressors, regulating valves, heat exchangers, and other apparatus including what is referred to as an ECS pack to condition engine bleed air. Modern day jet aircraft use turbocompressors in the ECS packs of their environmental controls systems to condition bleed air for use in the cabin wherein the turbocompressors are powered by the same bleed that is conditioned for cabin refreshing air and which is usually supplied by the gas turbine engines which provide aircraft propulsion. Other ECS systems employ auxiliary power units (APU's) that are separate gas turbine engines, not use for propulsion, to power ECS packs as well as other aircraft equipment.
Bleed air is conventionally taken from the engine compressor at a stage downstream of the variable vane compressor stages so as not to interfere with the operation of the variable vane stages which greatly enhance the efficiency of the gas turbine engine and greatly reduces the specific fuel consumption (SFC) of the engine. The compressor bleed air is cooled by fan air in a heat exchanger conventionally referred to as a precooler and is then delivered to the environmental control system for controlling cabin air freshness, pressure, and temperature. The ECS conventionally includes two or more ECS packs mounted in ECS bays on different sides of the aircraft which receive compressor bleed air from the engines. The bleed after being used to power the ECS pack and refresh the cabin is then dumped overboard. All the energy remaining in the bleed air dumped overboard cost fuel and therefore represents a significant loss in specific fuel consumption.
Extraction of aircraft bleed air from the engine compressor has adverse affects on the propulsion cycle and engine life. Engine turbine power is needed to compress air and account for compressor inefficiency. Therefore, extra fuel consumption is always associated with gas turbine engine compressor bleed air (air which does not produce thrust). This extra fuel burned in the engine combustor results in higher gas temperature delivered to the engine turbine and reduction of turbine blade life. Such penalties must be incurred in order that the engine turbine provide extra power associated with bleed air.
It is not possible, without undue complexity, to always bleed the engine compressor stage which provides exactly the correct pressure needed for the aircraft anti-ice and ECS systems. Typically, only two bleed ports are provided. Therefore, the result is to bleed air which exceeds minimum pressure requirements resulting in even higher penalty to the engine cycle than would be required by the aircraft systems.
Most often the bleed air is not only at a higher than required pressure, it is also too hot. For reasons of fire safety, maximum bleed air temperature is usually limited to 350.degree. to 500.degree. F. Temperature control requires cooling the bleed air with a precooler. Most modern engines use fan air to cool compressor bleed air. Use of fan air imposes an additional penalty on fuel consumption. Further, the precooler is usually large and requires a fan air scoop which produces drag. A typical large turbofan engine will consume about 2% extra fuel and run at about 20.degree. F. hotter turbine temperature in order to provide aircraft system bleed air. The present invention addresses these problems and deficiencies characteristic of the prior art and conventional apparatus used to supply aircraft bleed air.
FIG. 1 schematically illustrates an environmental control system (ECS) typical of the prior art having a conventional compressor bleed supply system 10 which extracts compressor bleed air from an aircraft propulsive gas turbine engine compressor section 8 to flow to and power a conventional ECS pack 30, which is depicted using an air cycle refrigeration system to cool and condition compressor bleed air, as is typical of the prior art. Compressor bleed supply system 10 includes a compressor mid-stage air bleed port 11 and a compressor discharge air bleed port 12 for supplying compressor bleed air through a compressor bleed air line 9. Bleed air normally flows through mid-stage bleed check valve 13 to shut-off valve 14. At low engine power, discharge bleed valve 15 can be opened, causing check valve 13 to close and bleed air to be delivered from compressor discharge port 12 to shut-off valve 14. Bleed air pressure is reduced to a duct structurally safe level by pressure regulator valve 16. Bleed air passes through engine bleed air precooler 17, which cools compressor bleed air, using cooler fan air from engine fan 19, to a safe temperature level before passing through duct 18 which is typically located in the aircraft wing near the aircraft fuel tanks. Fan cooling air from engine fan 19 flows through a precooler temperature control valve 21 upstream of precooler 17 and downstream of precooler 17 fan air is then flowed overboard of the engine as indicated by arrow 17a.
Having been partially reduced in both pressure and temperature, engine bleed air passes through duct 18 to ECS pack 30 for further temperature and pressure adjustment before introduction to aircraft 50 including its cabin, cockpit, and cargo bays as required. ECS flow control valve 31 drops bleed pressure substantially so that the pressure losses across ECS pack 30 maintains the desired bleed flow from the engine to the aircraft. The pressure losses across ECS pack 30 and in particular flow control valve 31 are very expensive in terms of fuel and thrust because of the energy spent compressing the bleed air.
A portion of the bleed flow passes through an ECS refrigeration compressor 32 to a ram air heat exchanger 33 then to an ECS refrigeration turbine 34. ECS refrigeration compressor 32 and ECS refrigeration turbine 34 are supplied as a single ECS turbocompressor assembly wherein ECS refrigeration compressor 32 is a centrifugal compressor and ECS refrigeration turbine 34 is a radial inflow turbine. The pressure drop across ECS refrigeration turbine 34 causes it to drive ECS refrigeration compressor 32.
Heat is removed from the bleed air by heat exchanger 33 which receives cooling air from a conventional ram air scoop 35a located on an outside surface of the aircraft and controlled by a ram air door 35. Ram cooling air is then dumped overboard as indicated by arrow 33a. Other portions of the bleed air bypasses ECS refrigeration compressor 32 through compressor bypass valve 36 or ECS pack bypass 37. ECS pack bypass air from valve 37 mixes with bleed air from ECS refrigeration turbine 34 for final temperature control of conditioned ECS bleed supply air that is supplied to aircraft 50 through an ECS air supply line 54.
Conditioned bleed air then flows to aircraft 50 for passenger flesh air consumption, cabin pressurization and temperature control. Temperature control is required to counter a varying cabin heat load 51 that includes aircraft skin cooling, solar heating, passenger body heating and electrical load heating. Part of cabin supply air 52 returns through re-circulation fans 53 and mixes with conditioned ECS bleed supply air from ECS air supply line 54. Cabin pressure is controlled by an outflow valve 55 that dumps overboard, as indicated by arrow 55a. Air flow from outflow valve 55 is essentially compressor bleed air taken from engine compressor section 8 and supplied to the aircraft.
Another problem addressed by the present invention relates to aerodynamic drag associated with engine nacelles, wings, pylons, tail sections and other aircraft outer surfaces. As air flows on to and over a surface such as an engine nacelle and aircraft wing it progressively builds up a low velocity boundary layer of increasing thickness. Within this boundary later a portion of the velocity component of free stream total pressure is converted to increased static pressure. As the result of rise in static pressure, boundary layer thickness, and diffusion a point is reached where back pressure causes an otherwise laminar boundary layer to become turbulent. In the turbulent region, a considerable amount of total pressure is converted to static temperature represented thermodynamically as an increase in entropy. By the time the boundary layer leaves the surface, or in the particular case of an aircraft gas turbine engine, the end of the nacelle, an unrecoverable loss in total pressure has occurred. The large entropy rise associated with turbulence is at the expense of air momentum. Turbulence also gives rise to increased static pressure which may increase the intensity of rearward acting pressure force on the surface. Now, if the boundary layer thickness is kept small, separation and turbulence will not occur or will be delayed and drag can be substantially reduced.
It is well known that one way to avoid increases or to reduce the build up in boundary thickness is to pump or bleed off boundary layer air through holes in the aircraft wetted surfaces such as the wing, tail, or portions of the engine nacelle. Boundary layer pumps or compressors would be desirable from an aerodynamic standpoint but, because of the relatively large airflow rates and added weight and complexity associated with effective boundary layer pumping or bleeding, the concept has not been adapted in modern aircraft and engines. Therefore, in one embodiment of the invention, this invention provides a means for effectively and economically using engine compressor bled air to power a nacelle boundary layer bleed compressor to bleed off laminar flow boundary layer air from the nacelle to reduce drag.
A similar problem was addressed in, and reference may be made to, patent application Ser. No. 07/531,718, invented by George A. Coffinberry, filed Jun. 1, 1990, and assigned to the same assignee and incorporated herein by reference.
Mechanically powered means for reducing boundary layer drag of various aircraft parts such as wings, nacelles, and aircraft tail assemblies have been proposed in the past and in patent application Ser. No. 07/489,150 entitled "AIRCRAFT ENGINE STARTER INTEGRATED BOUNDARY BLEED SYSTEM", invented by Samuel Davison, filed Mar. 6, 1990 and assigned to the same assignee. Other inventions addressing this problem are disclosed in a patent application Ser. No. 07/531,718 entitled "GAS TURBINE ENGINE POWERED AIRCRAFT ENVIRONMENTAL CONTROL SYSTEM AND BOUNDARY LAYER BLEED", filed on Jun. 1, 1990, and a patent application Ser. No. 07/572,825 entitled "AIRCRAFT ENGINE ELECTRICALLY POWERED BOUNDARY LAYER BLEED SYSTEM", filed on Jun. 1, 1990, both invented by the same inventor of this patent and assigned to the same assignee, both incorporated herein by reference.
Another patent application Ser. No. 07/531,734 entitled "GAS TURBINE ENGINE FAN DUCT BASE PRESSURE DRAG REDUCTION", invented by the same inventor of this patent, filed on Jun. 1, 1990, and assigned to the same assignee, and incorporated herein by reference, discloses a mechanically powered means of bleeding boundary layer air and reducing the drag of the aircraft by introducing at least a portion of the pressurized bleed air into the fan duct of the engine to reduce the base drag of the duct.